1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine airfoil.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98.
In a gas turbine engine, air is first compressed to a high pressure in a compressor. The high pressure air is then mixed with fuel and burned at nearly constant pressure in the combustor. The high temperature gas exhausted from the combustor is then expanded through a turbine which then drives the compressor. If executed correctly, the exhaust stream from the turbine maintains sufficient energy to provide useful work by forming a jet, such as in aircraft jet propulsion or through expansion in another turbine which may then be used to drive a generator like those used in electrical power generation. The efficiency and power output from these machines will depend on many factors including the size, pressure and temperature levels achieved and an agglomeration of the efficiency levels achieved by each of the individual components.
Current turbine components are cooled by circulating relatively (to the gas turbine engine) cool air, which is extracted from the compressor, within passages located inside the component to provide a convective cooling effect. In many recent arrangements, the spent cooling flow is discharged onto the surfaces of the component to provide an additional film cooling effect.
The challenge to cool first stage turbine vanes (these are exposed to the highest temperature gas flow), in particular, is complicated by the fact that the pressure differential between the vane cooling air and the hot gas which flows around the airfoil must necessarily be small to achieve high efficiency. Specifically, coolant for the first stage turbine vane is derived from the compressor discharge, while the hot gas is derived from the combustor exit flow stream. The pressure differential available for cooling is then defined by the extremely small pressure drop which occurs in the combustor. This is because the pressure of the coolant supplied to the vane is only marginally higher than the pressure of the hot gas flowing around the airfoil as defined by the combustor pressure loss, which is desirably small. This pressure drop is commonly on the order of only a few percentage points. Further, it is desirable to maintain coolant pressure inside the vane higher than the pressure in the hot gas flow path to insure coolant will always flow out of the vane and not into the vane. Conversely, in the event hot gas is permitted to flow into the vane, serious material damage can result as the materials are heated beyond their capabilities and progression to failure will be swift. As a consequence, current first stage turbine vanes are typically cooled using a combination of internal convection heat transfer using single impingement at very low pressure ratio, while spent coolant is ejected onto the airfoil surface to provide film cooling. FIG. 6 shows an example of this prior art turbine airfoil cooling circuit.
The efficiency of the convective cooling system is measured by the amount of coolant heat-up divided by the theoretical heat-up possible. In the limits, little coolant heat-up reflects low cooling efficiency while heating the coolant to the temperature of the surface to be cooled (a theoretical maximum) yields 100% cooling efficiency. In the previous methods using single impingement, the flow could only be used once to impinge on the surface to be cooled. This restriction precludes the ability to heat the coolant substantially, thereby limiting the cooling efficiency.
BRIEF SUMMARY OF THE INVENTION
It is an object of the present invention to provide for a turbine airfoil with improved cooling efficiency over the cited prior art turbine airfoils.
It is another object of the present invention to provide for an air cooled turbine airfoil with internal cooling air passages that cannot be formed from the investment casting process.
It is another object of the present invention to provide for an air cooled turbine airfoil of the spar and shell construction.
It is another object of the present invention to provide for an air cooled turbine airfoil having a thin near-wall cooled airfoil surface.
These objectives and more are achieved with the air cooled turbine airfoil that is made up of a number of stacked laminates each having a specific cooling air circuit with each layer separated by a divider layer in order that a sequential impingement cooling passage is formed to provide impingement cooling of the airfoil. In one embodiment, the sequential impingement cooling circuit is formed using three different laminates stacked with a divider layer separating each of the other two layers and forms a cooling circuit that provides a first impingement cooling of the forward half of the pressure side airfoil wall, then a second impingement cooling of the aft half of the pressure side wall, and then a third impingement cooling of the entire suction side wall of the airfoil before discharging the spent cooling air through a row of exit cooling holes in the trailing edge of the airfoil. The alternating layers for a spar for the turbine blade or vane which forms a structural support for a shell that forms the airfoil shape of the blade or vane.
In a second embodiment of the present invention, the air cooled airfoil with the sequential cooling circuit described above is formed in a single piece airfoil by a metallic or ceramic printing process that can form the entire airfoil on a molecular level.
This printing process is developed by Mikro Systems, Inc. of Charlottesville, Va.